The assignee of the present invention designs and manufactures spacecraft that include various structures that are deployable from a launch configuration to an on-orbit, or deployed configuration. For example, it is known to reconfigure a solar array from a launch configuration to an on-orbit configuration. In order to achieve compatibility with launch vehicle fairing constraints, the solar array, in the launch configuration, is generally disposed as compactly as possible adjacent to a sidewall of the spacecraft body. In the on-orbit configuration, the solar array is generally deployed to extend a very substantial distance (several times the width of the spacecraft body) along an axis (the “a axis”) that is substantially orthogonal to the sidewall. During orbital operation, the solar array may be caused to rotate about the α axis so as to maximize collection of solar radiation by photovoltaic solar cells disposed on a surface of the solar array. A spacecraft solar array generally includes a closed cable loop (CCL) system which synchronizes deployment of various portions of the solar array. In the absence of the presently disclosed techniques, such a CCL system prevents independent articulation of solar arrays about an axis (the “β axis”) that may be transverse to, or at least non-aligned with, the α axis. As a result, a long axis of the solar array, once deployed in the on-orbit configuration, coincides with the α axis and the solar array is latched in that position for the entire mission.